1. Technical Field
This invention relates to gas turbine engine stator vanes and rotor blades in general, and to stator vanes and rotor blades possessing internal cooling apparatus in particular.
2. Background Information
In the turbine section of a gas turbine engine, core gas travels through a plurality of stator vane and rotor blade stages. Each stator vane or rotor blade has an airfoil with one or more internal cavities surrounded by an external wall. The suction and pressure sides of the external wall extend between the leading and trailing edges of the airfoil. Stator vane airfoils extend spanwise between inner and outer platforms and the rotor blade airfoils extend spanwise between a platform and a blade tip.
High temperature core gas (which includes air and combustion products) encountering the leading edge of an airfoil will diverge around the suction and pressure sides of the airfoil, or impinge on the leading edge. The point along the leading edge where the velocity of the core gas flow goes to zero (i.e., the impingement point) is referred to as the stagnation point. There is a stagnation point at every spanwise position along the leading edge of the airfoil, and collectively those points are referred to as the stagnation line. Air impinging on the leading edge of the airfoil is subsequently diverted around either side of the airfoil.
Cooling air, typically bled off of a compressor stage at a temperature lower and pressure higher than the core gas passing through the turbine section, is used to cool the airfoils. The cooler compressor air provides the medium for heat transfer and the difference in pressure provides the energy required to pass the cooling air through the stator or rotor stage. Film cooling and internal convective/impingement cooling are prevalent airfoil cooling methods. Film cooling involves cooling air bled from an internal cavity which forms into a film traveling along an exterior surface of the stator or rotor airfoil. The film of cooling air increases the uniformity of the cooling and insulates the airfoil from the passing hot core gas. A person of skill in the art will recognize, however, that film cooling is difficult to establish and maintain in the turbulent environment of a gas turbine.
Convective cooling, on the other hand, typically includes passing cooling air through a serpentine of passages which include heat transfer surfaces such as "pins" and "fins" to increase heat transfer from the airfoil to the cooling air passing therethrough. Convective cooling also typically includes impingement cooling wherein cooling air jets through a metering hole, subsequently impinging on a wall surface to be cooled. An advantage of impingement cooling is that it provides localized cooling in the impinged upon region, and can be selectively applied to achieve a desirable result. A disadvantage of impingement cooling is that the convective cooling provided by the impingement is limited to a relatively small surface area. As a result, a large number of cooling apertures are required to cooling extended areas.
What is needed, therefore, is an airfoil with an internal cooling scheme that provides cooling more efficiently than is possible with presently available airfoils, one that promotes film cooling along the outside of the airfoil's exterior wall, and one that can be readily manufactured.